It is known that gas turbines have components in the turbine or combustion chamber which are exposed to impingement by hot gas and which are cooled by means of cooling air. Cooling of the components is carried out so that these permanently withstand the hot temperatures of the hot gas. Compressor air, which can be bled from the compressor main flow at different points of the compressor, is used as cooling air, as is known. For this, FIG. 3 shows in a longitudinal section a casing-side cooling air bleed in a compressor 10 of a gas turbine 12. The bleed of compressed air 42 from the air 18 which flows in an annular flow path 14 is carried out via a row of radially disposed bleed openings 40 which are arranged in the outer wall 26 which delimits the flow path 14 on the outside. All the bleed openings 40 lead to an annular chamber 44 which is arranged outside the outer wall 26 and in which the extracted air is collected. The chamber 44 itself is encompassed, and therefore delimited, radially on the outside by the casing 32 of the compressor 10. In the casing 32, again a small number of larger openings 64, for example three or four openings, are distributed over the circumference, from which the part 42 of compressed air which is fed to the annular chamber 44 can be discharged in order to be fed via further pipes, which extend outside the gas turbine and are not additionally shown, to the turbine or to the combustion chamber where the extracted air is used as specified for the purpose of cooling the components which are exposed to impingement by hot gas.
In the annular flow path 14 of the compressor 10, stator blades 20 and rotor blades 22 are arranged in an alternating manner in rings in each case, which stator blades 20 are fastened in a rotationally fixed manner on the casing side, i.e. on the outer wall 26, and which rotor blades 22 are fastened on the inner wall 24 of the annular flow path 14, i.e. on the rotor 24 and rotatable with this. The tips of the rotor blades 22 lie opposite the outer wall of the flow path 14, forming a gap 36, just as the tips of unshrouded stator blades 20 lie opposite the inner wall 24 of the flow path 14. In order to keep these gaps between rotor 24 and stator as small as possible for each operating phase, it is necessary to thermally balance the rotor 24 and the stator as efficiently as possible.
Depending upon the system, the rotors, on account of their large mass and on account of the smaller inner throughflow, are thermally slower than the casing and expand correspondingly slower. In particular, the radial and axial expansions which occur during transient operating conditions, for example during cold starting or during shutdown of the gas turbine, have to be taken into consideration in the design of the necessary radial and axial gaps. This has a disadvantageous effect upon the required installation space, upon the robustness of the arrangement and particularly upon the gas turbine efficiency overall.
Furthermore, an axial compressor for an aircraft gas turbine with an air bleed is known from U.S. Pat. No. 5,203,162. Upstream of the air bleed opening, which is arranged in the casing, there are two radially staggered annular chambers which are separated from each other by means of a common partition. In this case, the inner delimiting wall of the inner annular chamber is formed by the annular wall which forms the flow path of the compressor. Provision is made both in the annular wall and in the partition for openings in this case, through which the partial flow which is extracted from the compressor can flow out. The openings of the annular wall and of the partition are arranged in this case in an axially offset manner in relation to each other in order to trap and to deflect the particles which are entrained by the main flow and which originate for example as a result of abrasion by the compressor rotor blades on the casing, since impingement of the particles upon the comparatively thin outer casing wall is to be avoided.
Furthermore, a compressor with a plurality of bleed openings is known from U.S. Pat. No. 5,160,241, in which the air flow which is extracted from the compressor main flow is routed via an annular chamber. For avoiding varying thermal load along the circumference, provision is made at each bleed opening on the collecting chamber side for deflection plates which extend in a sector-like manner in the circumferential direction. The guide plates in this case are provided only locally in the region of the bleed openings in each case. By means of the deflection plates, the air which flows into the collecting chamber being able to flow along the inner surface of the casing is avoided, as a result of which the unequal thermal load of the casing can be reduced.
Moreover, a guide ring segment of a turbine is known from U.S. Pat. No. 4,303,371, which by means of a U-shaped plate which is arranged therein can be impingement-cooled.